The invention relates to reducing base pressure drag of aircraft. More particularly the invention relates to reducing base pressure drag in low pressure regions of aircraft gas turbine engine fan ducts.
Means for reducing boundary layer drag of various aircraft parts such as wings, nacelles, and aircraft tail assemblies have been proposed in the past and in patent application Ser. No. 07/489,150 entitled "AIRCRAFT ENGINE STARTER INTEGRATED BOUNDARY BLEED SYSTEM", invented by Samuel Davison, filed Mar. 6, 1990 and assigned to the same assignee and in a patent application Ser. No. 071,531,718 entitled "GAS TURBINE ENGINE POWERED AIRCRAFT ENVIRONMENTAL CONTROL SYSTEM AND BOUNDARY LAYER BLEED", invented by the same inventor of this patent, filed on Jun. 1, 1990, and assigned to the same assignee, both incorporated herein by reference. One of the advantages of those inventions is the reduction in aerodynamic drag associated with engine nacelles, wings, pylons, tail sections and other aircraft outer surfaces due to boundary layers going turbulent. As air flows on to and over a surface such as an engine nacelle and aircraft wing it progressively builds up a low velocity boundary layer of increasing thickness. Within this boundary layer a portion of the velocity component of free stream total pressure is converted to increased static pressure. As the result of rise in static pressure, boundary layer thickness, and diffusion a point is reached where back pressure causes an otherwise laminar boundary layer to become turbulent.
In the turbulent region, a considerable amount of total pressure is converted to static temperature represented thermodynamically as an increase in entropy. By the time the boundary layer leaves the surface, or in the particular case of an aircraft gas turbine engine the end of the nacelle, an unrecoverable loss in total pressure has occurred. The large entropy rise associated with turbulence is at the expense of air momentum. Turbulence also gives rise to increased static pressure which may increase the intensity of rearward acting pressure force on the surface. Now, if the boundary layer thickness is kept small, separation and turbulence will not occur or will be delayed and drag can be substantially reduced.
One way to avoid increase in boundary thickness is to pump or bleed off boundary layer air through holes in the surface. This however incurs a ram air drag penalty due to loss of momentum of the boundary layer air taken on board the engine. Engine net thrust is equal to engine exhaust momentum minus inlet ram drag. Therefore the result is a problem of decreased engine efficiency or increased engine specific fuel consumption. Another problem addressed by the present invention relates to base pressure drag in the aircraft gas turbine engine fan duct. Certain low pressure regions of the fan duct such as in the aft tapered airfoil section of a pylon fairing produces base pressure drag. A description of low pressure base pressure drag may be found in the classical reference on the subject of drag, FLUID-DYNAMIC DRAG by Sighard Hoerner (particularly pages 3-18 to 3-20 and page 20-16), which is incorporated herein by reference. Thus, the present invention provides a means for ducting air brought on board an aircraft and compressed to a low pressure area of a gas turbine engine fan duct which otherwise produces base pressure drag in the duct.
It is, therefore, an object of the present invention to provide a more efficient aircraft gas turbine engine by using air that has been brought on board and thereby producing ram air drag, to reduce base pressure drag of the aircraft.
Another object of the present invention to provide a more efficient aircraft gas turbine engine by using compressed air that has been brought on board an aircraft gas turbine engine, thereby producing ram air drag, to reduce base pressure drag in the aircraft engine fan duct.
These objects and other features and advantages will become more readily apparent in the following description when taken in conjunction with the appended drawings.